ANTENNA TECHNOLOGY FOR QUASAT APPLICATION

John S . Archer and W i l l i a m B. Palmer TRW Elec t ronics & Defense Sec tor

Redondo Beach, Ca l i fo rn ia

Large Space Antenna Systems Technology - 1984 December 4 - 6 , 1984

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https://ntrs.nasa.gov/search.jsp?R=19850015517 2020-03-23T16:06:11+00:00Z

INTRODUCTION

This paper summarizes the results of a TRW study performed to provide JPL with antenna cost and design data for use in assessing the technical feasibility of QUASAT, a very long baseline interferometry (VLBI) mission in space. Science requirements for this mission have been defined by the NASA OVLBI Technical Working Group and the European Space Agency QUASAT Working Group (Paris, Oct. 1983). The mission concept involves a free-flying high-earth-orbiting spacecraft which carries a space-deployable radio antenna of about 15 meters in diameter with a receiving capa- bility at K, C, and L bands.

TRW's approach to the requirements of the QUASAT antenna configuration adapted a hybrid growth version of the advanced Sunflower, or precision deployable, antenna concept. The basic precision deployable antenna concept uses a furlable rigid-panel reflector design as illustrated in figure 1, where it is shown mounted in the Shuttle bay.

Figure 1. Precision deployable antenna in the Shuttle's bay.

252

PRECISION DEPLOYABLE CONCEPT

TRW has been actively investigating the design of precision deployable antennas for several years. Figure 2 is an illustration of a 12-inch-diameter working model. Developed for a precision-contour reflector application, the concept can accommodate Cassegrain or focal-fed antenna applications requiring high-efficiency high-frequency apertures.

Figure 2. Deployable symmetrical aperture.

253

27.5-FOOT RESEARCH MODEL

F igures 3 and 4 show a des ign and t h e t o o l i n g which have been developed i n TRW's ongoing r e s e a r c h and development program f o r t h e p r e c i s i o n deployable antenna. Information from t h i s ongoing e f € o r t w a s used t o p r e p a r e t h e technology and c o s t i n g d a t a provided i n t h e f i g u r e s which fol low.

To provide a s t i f f , near-continuous s u r f a c e , t h e antenna s e c t i o n s c o n s i s t of r i g i d pane ls of graphite-epoxy f a c e s h e e t s cover ing aluminum honeycomb c o r e sandwich. The c e n t e r s e c t i o n , t h e one-piece honeycomb sandwich s t r u c t u r e , i s c e n t e r e d w i t h i n t h e f o l d i n g sandwich s e c t i o n s .

The s i x main f o l d i n g p a n e l s are hinged t o a c a n t i l e v e r e d suppor t r i n g a t t a c h e d t o t h e p e r i p h e r y of t h e c e n t e r s e c t i o n . Between t h e s i x main panels on t h e c a n t i - l e v e r e d r i n g are s i x p a i r s of i n t e r m e d i a t e panels . Two o r more h inges a long a d j a c e n t edges connect each p a i r of i n t e r m e d i a t e pane ls t o t h e main panels and t o each o t h e r . The h inges have a d j u s t a b l e s t o p s which l o c a t e t h e pane l s u r f a c e s a c c u r a t e l y i n t h e deployed c o n f i g u r a t i o n . Spr ings i n t h e h inges d r i v e t h e p a n e l s t o t h e deployed pos i - t i o n . To ensure s y n c h r o n i z a t i o n of a l l pane ls d u r i n g deployment, a compound univer- s a l coupl ing i n t e r c o n n e c t s a d j a c e n t inboard h inges of t h e main panels . E i t h e r a damping device o r a geared motor c o n t r o l s t h e deployment rate. On ordnance-actuated p i n p u l l e r supported on one of t h e tie-down f i t t i n g s r e s t r a i n s t h e f u r l e d antenna. The number of p a n e l s hinged from the f i x e d c e n t e r s e c t i o n can be v a r i e d as r e q u i r e d t o opt imize t h e d iameter and h e i g h t of t h e stowed antenna.

F i g u r e 3. 27.5-foot-diameter deployable s o l i d r e f l e c t o r (stowed).

I

254

Figure 4 . Master layup mold.

255

QUASAT REQUIREMENTS

The following specific mechanical requirements constrained the TRW QUASAT study:

( 1 ) The antenna shall be a deployable reflector configuration.

( 2 ) The antenna shall be stowable within the Shuttle Orbiter bay.

(3) The antenna characteristics shall be assessed for a range of apertures from 10 (33) to 20 ( 6 6 ) meters (feet).

( 4 ) The antenna surface error shall not exceed 0.8 millimeter (0.031 inch).

(5) The mechanical parameters to be assessed shall include antenna weight, mass properties, stowed envelope, structural interface details, manufactured surface contour, thermal distortion, and materials of construction.

( 6 ) Cost and schedule parametric data shall be provided.

I 256

QUASAT HYBRID ANTENNA CONFIGUATION

Technical paramet r ic d a t a a r e provided f o r antenna diameters of 10 (331, 15 (491, and 20 (66) meters ( f e e t ) . For each antenna diameter ( D ) , we s e l e c t e d a f o c a l length (F) t o maintain a cons tan t F/D r a t i o of 0.4. I n s e l e c t i n g the antenna con- f i g u r a t i o n s f o r each diameter , we designed t h e stowed envelope dimensions t o f i t wi th in the S h u t t l e launch veh ic l e envelope. This c o n s t r a i n t l i m i t s the maximum al lowable p rec i s ion Sunflower ape r tu re t o 33 f e e t . A 33-foot-diameter ape r tu re stows i n a c y l i n d r i c a l envelope 14.5 f e e t i n diameter by 12.2 f e e t long. Any l a r g e r diam- eter would exceed t h e a v a i l a b l e S h u t t l e envelope diameter. The hybrid approach, which we d i scuss below, uses t h e p rec i s ion deployable r e f l e c t o r as the core of a l a r g e r ape r tu re . The hybrid concept w a s s e l e c t e d because i t o f f e r s a p r a c t i c a l means of e f f e c t i v e l y extending the deployable concept wi th i t s p rec i s ion contour advantage t o diameters l a r g e r than 33 f e e t while remaining wi th in a stowed envelope of 14.5- foo t maximum diameter.

The conf igu ra t ions s e l e c t e d are based upon a common, r ig id-panel , p rec i s ion deployable r e f l e c t o r with an ape r tu re diameter of 33 f e e t , a s i z e which remains con- s t a n t f o r a l l antenna diameters. Such a c e n t r a l deployable r e f l e c t o r i s i l l u s t r a t e d i n f i g u r e s 5 t o 7. The r e f l e c t o r has the inherent c a p a b i l i t y of being manufactured wi th a very low contour roo t mean square (rms), i n t h e range of 0.005 t o 0.015 inch. For t h e purpose of t h i s s tudy , a value of 0.014 inch was chosen as s u f f i c i e n t t o s a t i s f y t h e s y s t e m requirements.

A s t h e a p e r t u r e diameter i nc reases over 33 f e e t , t h e diameter of the p rec i s ion deployable po r t ion holds cons tan t a t 33 f e e t . The inc rease i n t o t a l ape r tu re diam- e t e r beyond 33 f e e t i s a t t a i n e d by c a n t i l e v e r i n g an annular d i s k s t r u c t u r e from the 33-foot-diameter p r e c i s i o n deployable s t r u c t u r e . The na ture of t h e annular d i s k s t r u c t u r e and i t s attachment t o t h e 33-foot r e f l e c t o r is such t h a t both the annular d i sk and t h e suppor t ing c e n t r a l s t r u c t u r e deploy s imultaneously, as a s i n g l e s t r u c - t u r a l system.

The behavior of t h e annular d i sk s t r u c t u r e i s t h a t of a f l e x i b l e membrane. The advantages of t h i s cons t ruc t ion over a c e n t r a l r igid-panel s t r u c t u r e are l i g h t e r weight and f l e x i b i l i t y , a l lowing f o r storage within the Shuttle envelope, as seen in f i g u r e 7. The disadvantage is h igher contour rms to l e rances compared t o t h e c e n t r a l 33-foot por t ion .

A 49-foot antenna of hybrid concept c o n s i s t s of a 33-foot-diameter cen te r s t ruc - t u r e surrounded by an annular d i sk of 33 f e e t i n s i d e diameter (i .d.1 and 49 f e e t o u t s i d e diameter (o.d.1, t h e 33-foot cen te r s t r u c t u r e being a p rec i s ion deployable r e f l e c t o r . The 33/49-foot d i s k with a f l e x i b l e membrane su r face s t r u c t u r a l l y can t i - l e v e r s from t h e c e n t r a l p rec i s ion deployable r e f l e c t o r . S imi l a r ly , a 66-foot antenna of t h i s hybrid concept c o n s i s t s of a 33-foot cen te r s t r u c t u r e surrounded by an annu- l a r d i s k of 33 f e e t i.d. and 66 f e e t 0.d. Again, t h e 33-foot cen te r s t r u c t u r e i s a p r e c i s i o n deployable r e f l e c t o r , and the 33166-foot d i sk s t r u c t u r e i s a f l e x i b l e mem- brane s u r f a c e s t r u c t u r a l l y can t i l eve red from the c e n t r a l p r e c i s i o n deployable r e f l e c t o r .

257

QUASAT HYBRID ANTENNA CONFIGURATION ( CONCLUDED)

I Constraint and Release

The stowed conf igura t ion f o r t he 33-foot ape r tu re i s i l l u s t r a t e d i n f i g u r e 5. Extension arms from t h e in te rmedia te p e t a l s are locked toge the r i n the stowed case t o provide a r i g i d i n t e g r a t e d s t r u c t u r e f o r r e a c t i n g t h e launch loads. For the 49- and 66-foot ape r tu re s with the can t i l eve red annular d i sk s t r u c t u r e s , an e x t e r n a l r i n g , when stowed, r i g i d l y cons t r a ins the c e n t r a l deployable s t r u c t u r e , as i l l u s t r a t e d i n f i g u r e 6 . The ou te r edges of t he annular d i sk are cons t ra ined a t a poin t above t h e feed support s t r u c t u r e . The release mechanism is shown i n V i e w A-A. In both cases , r e l e a s e i s e f f e c t e d by pyrotechnic p in p u l l e r s , and deployment i s ac tua ted by pass ive sp r ings and con t ro l l ed by hydraul ic dampers.

I Support I n t e r f a c e

The support i n t e r f a c e attachment t o the spacec ra f t i s i l l u s t r a t e d i n f i g u r e s 5 and 6 , and the f i t t i n g loca t ions a r e i d e n t i f i e d i n f i g u r e 7. This i n t e r f a c e is identical for all apertures .

Materials

The antenna i s f a b r i c a t e d of graphite-epoxy materials. The c e n t r a l deployable r e f l e c t o r is cons t ruc ted of sandwich panels made of g r a p h i t e f i b e r r e in fo rced p l a s t i c (GFRP) f aceshee t s over an aluminum honeycomb core. Deep r i b s of t h e same materials r e i n f o r c e t h e panels . These materials provide a r i g i d , thermally s t a b l e s t r u c t u r e of r e l a t i v e l y l i g h t weight. The antenna su r face i s coated wi th vacuum-deposited a lu- minum t o provide high RF r e f l e c t i v i t y . The f i n a l coa t ing is of a 0.003- t o 0.004- inch-thick thermal c o n t r o l pa in t which minimizes s p e c u l a r i t y and reduces the equi- l i b r ium temperature of t he r e f l e c t o r s t r u c t u r e i n the o r b i t a l s o l a r environment.

LD 4LL DEPLOY IIYVLTANEOUSLY

Figure 5. 33-foot-diameter antenna.

258

/

PIN PULLERS PLACES

Figure 6. Configuration of 49- and 66-foot-diameter antennas.

147”

1 33-foot DIAMETER 49-foot DIAMETER 66-foot DIAMETER

3 PL ON 60” RADIUS

VIEW A-A LOCATION OF SPACECRAFT INTERFACE FllTlNGS

Figure 7. Stowed envelopes (F/D = 0.4) .

2 59

QUASAT FLIGHT CONFIGURATION

The QUASAT flight configuration using the hybrid antenna configuration is illus- trated in figure 8. System (STS) bay is illustrated in figure 9.

The furled configuration stowed in the Space Transportation

R N R F S ELECTRONICS .

MAGNETOMETER

Figure 8. QUASAT flight configuration for advanced sunflower reflector.

ncs THRUSTER roD

Figure 9. QUASAT advanced sunflower antenna stowed for launch in STS bay.

260

HYBRID ANTENNA WEIGHT

D IlnI)

33

48

88

TRW's experience with s o l i d r e f l e c t o r cons t ruc t ion of many types has shown t h a t antenna weights vary from 0.4 t o 1.0 psf of t he p ro jec t ed aper ture . value i s from a very conserva t ive design, which is about 15 years o ld , whereas t h e lowest value is i n d i c a t i v e of cu r ren t l igh tweight l a r g e r e f l e c t o r experience. To represent t he c e n t e r po r t ion of t he antenna, we have used an average value of 0.6 t o 0.7 psf inc luding t h e weight of t he feed support tower and launch c o n s t r a i n t mechanisms.

The h ighes t

WEIGHT IIbl

CENTER ANNULAR

STRUCTURE STRUCTURE DEPLOYABLE DISK TOTAL

sar 0 500

550 110 (IBO

Ba) 2% 855

The weight of t h e annular d i sk s t r u c t u r e i s es t imated a t 0.1 psf. The t o t a l r e f l e c t o r weight and the breakdown i n weight between the cen te r deployable s t r u c t u r e and t h e o u t e r d i s k are t abu la t ed i n f i g u r e 10 and p l o t t e d i n of 33, 4 9 , and 66 f e e t .

f i g u r e 11 f o r ape r tu re s

Figure 10. Weight d i s t r i b u t i o n i n antenna s t r u c t u r e .

WEIGHT Ilbl

33 48 88

APERTURE Itrail

Figure 11. Antenna weight versus ape r tu re .

261

HYBRID ANTENNA MASS PROPERTIES

The center of gravity (cg) and the moments of inertia have been based on the stowed and deployed geometric envelopes illustrated previously. The antenna focal- length-to-diameter ratio is constant at 0.4 €or all apertures from 33 to 66 feet. Furthermore, the antenna mass is largely concentrated in the central portion of the structure. This results in a relatively constant value of the longitudinal position of the cg for the deployed antenna, except €or the effect of the feed support tower.

In the stowed configuration, the cg position estimate is based on the stowed envelope geometry and the mass distribution whereas the estimate of the cg position in the deployed case is based on the deployed envelope geometry and the mass distri- bution. Figure 12 provides the calculated cg data.

I The roll moments of inertia when stowed and deployed are provided in figures 13 and 14.

(DISTANCE FORWARD

OF VERTEX) CG

(feet)

.+------- 4+ DEPLOYED 'm

33 49

APERTURE (feet1

66

Figure 12. Center of gravity versus aperture.

262

33 49 66

APERTURE (feet)

F i g u r e 13. R o l l moment of i n e r t i a v e r s u s a p e r t u r e ( s towed) .

JD (1s Ib - ft2)

33 49

APERTURE (feet)

66

F i g u r e 14. R o l l moment of i n e r t i a v e r s u s a p e r t u r e (dep loyed) .

263

HYBRID ANTENNA FREQUENCY

The fundamental frequency of the proposed concept at the various apertures of 33 , 4 9 , and 66 feet is based on extrapolation from a known configuration. A detailed vibration analysis was performed on a 20-foot-diameter precision deployable reflector aperture. The fundamental frequency in the stowed configuration was 12 Hz and in the deployed configuration, 5 Hz. apertures being studied by using a single-degree-of-freedom model and accounting for mass and size changes. Figure 15 depicts the results of our analysis.

These data have been extrapolated to the

1

33 49 66 APERTURE (feet)

Figure 15. Fundamental frequency versus aperture.

6

5

3

2

I 264

HYBRID ANTENNA CONTOUR TOLERANCE

The hybrid antenna contour distortion is defined as rms deviation from a perfect parabola. The two principal sources of contour error are manufacturing tolerances and thermal distortion on orbit. The total manufacturing error occurs in three areas - fabrication, assembly, and deployment repeatability.

The basic fabrication error is incurred when making the individual panels of the central deployable structure of the antenna. This error, based on experience with different materials and designs, is shown in figure 16. The proposed construction of graphite-epoxy, rib-supported shells will provide a contour rms-to-diameter ratio of 0.00003. The largest panel of the central reflector has a maximum dimension of 12 feet. We may, therefore, expect the panels to individually have an rms of 0.004 inch under ideal conditions. To be conservative, we allowed a 100-percent safety factor and, in addition, assumed that assembly and repeatability errors have the same order of magnitude. The net result is a total rms, including all effects, of 1.732 times 2 times 0.004 = 0.014 inch for the central 33-foot portion of the antenna.

The rms of the annular disk portion of the antenna is related to the method of support of the flexible membrane. The support structure is selected to maintain a fabricated rrns of 0.033 inch for the 66-foot disk. The assembly and deployment repeatability errors are assumed to equal those for the central deployable structure.

The net results for the center and annular segments and for the complete 33-, 49- , and 66-foot antennas are tabulated in figure 17. The total combined rms is determined by root-sum-squares using area weighting for the components.

We estimated the thermal distortion by extrapolation from existing analyses on smaller apertures. A conservative value for graphite-epoxy construction is 0.001-inch rms for a 5-foot-diameter aperture. Linear extrapolation gives rms values of 0.007 inch for 33 feet, 0.010 inch for 49 feet, and 0.013 inch for 66 feet. These values, found in figure 17, when combined with the net manufactured rms give the total on-orbit rrns values for each case.

The total on-orbit rms performance satisfies the system requirements for all aperture sizes.

NORMALIZED CONTOUR RMS (6lD) MATERIAL CONFIGURATION

FIB€ RGLASS SANDWICH SHELL > 5 X lo4 ADJUSTED SHELL 0.7 X lo4

0.6 X lo4 KEVLAR RIBSUPPORTED SHELL

ALUMINUM SANDWICH SHELL 0.5 X lo4

GRAPHITE SANDWICH SHELL 0.g x 10-4 0.3 X lo4 RIB-SUPPORTED SHELL

ADJUSTED SHELL 0.14 X lo4

Figure 16. Nondeployable reflector performance correlation.

265

33-49ft 49ft

11.25) (2.25)

49-66ft 66ft

(1.75) (4)

NET MANUFAG 1 0.014 1 0.028 1 0.023 1 0.035 1 0.029 TURED RMS

FABRICATION 0.008

ASSEMBLY 0.008

DEPLOYMENT 0.008

0.025 0.033

0.008 0.008

\ 0.008 0.008 \

THERMAL DISTRIBUTION

Figure 17 . Contour RMS versus aperture.

0.007 0.010 0.013

266

0.025

~

TOTAL 0.032

I MODIFIED 49-FOOT STOWED CONFIGURATION

STS launch costs are strongly dependent upon the length of bay occupied by the The principal problem with the hybrid configura- payload to be carried into orbit.

tion is the length of the stowed antenna envelope.

The stowed envelope can be shortened by almost 8 feet for the 49-foot antenna by folding the outer-rib mesh annular disk onto the surface of the solid segmented panels of the inner 33-foot-diameter precision deployable antenna. This is accom- plished by changing the hinge mechanism on the outer ribs and revising the stowage sequence. The solid panels of the precision deployable antenna are subsequently stowed while supporting the folded ribs of the outer annular portion. This stowage technique is illustrated in figures 18 to 20. It cannot be used for the 66-foot antenna because the ribs in the annular disk are too long to be folded onto the folding panels of the 33-foot central antenna structure.

In order to take full advantage of the reduced length of the STS bay required for launch, the feed support structure must also be made stowable, as shown in figure 21, which illustrates the revised stowage envelope in the STS bay.

Figure 18. Deployed 49-foot hybrid antenna with modified stowage technique.

267

Figure 19. Partially furled 49-foot hybrid antenna with modified stowage technique.

Figure 20. Completely furled 49-foot hybrid antenna with modified stowage technique.

I 268

EXTENDABLE FEED SUPPORT TRUSS,

PRECISION DEPLOYABLE SCIENCE ANTENNA

NSTS

RCS THRUSTER POD

NSTS KEEL VF /

Figure 21. QUASAT 49-foot hybrid antenna stowed in STS bay for launch .with modified stowage envelope.

2 69

CONCLUSION

I

For the QUASAT antenna, the hybrid concept discussed here is a promising altern- ative to an all-mesh contour. The greater efficiency of the solid antenna surface, the tighter tolerances possible with the solid folding panels, and the potential for using the center 33-foot portion for higher RF frequencies than possible with mesh surfaces open up interesting prospects for future accomplishments. The possibility of storing a 49-foot antenna within a short envelope in the Shuttle's bay contributes immensely to the feasibility of the concept.